CN112486196B - Aircraft rapid trajectory optimization method meeting strict time and position constraints - Google Patents

Aircraft rapid trajectory optimization method meeting strict time and position constraints Download PDF

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CN112486196B
CN112486196B CN202011392656.4A CN202011392656A CN112486196B CN 112486196 B CN112486196 B CN 112486196B CN 202011392656 A CN202011392656 A CN 202011392656A CN 112486196 B CN112486196 B CN 112486196B
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韦常柱
李瑜
佘智勇
樊雅卓
乔鸿
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Harbin Institute of Technology
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Abstract

Hair brushThe invention discloses a rapid track optimization method of an aircraft, which meets strict time and position constraints. Step 1: setting parameters; the quasi-state parameters include load at t1Time on track, standard on track point r1(ii) a Assuming that the load is t through the online re-planning and the self-adaptive guidance of the track2Time on track, actual point on track is r2(ii) a Step 2: defining a point coordinate system; origin O of coordinate systemPIs the center of the earth, xpThe shaft is positioned on a connecting line of the geocentric and the near point of the target track and points to the near point; and step 3: based on the parameters and point coordinate systems in the step 1 and the step 2, the concept of the approximate point angle phi is utilized to calculate the aircraft slave r1Fly to r2Time Δ t of; and 4, step 4: and (4) correcting the satellite orbit entering time deviation by using the step 1-3 and a core secondary boot time iterative correction method. The method is used for solving the problems that the thrust of a high-thrust liquid rocket engine applied to a carrier rocket can not be adjusted, and the entry point can not be accurately controlled, namely the entry position can not be restrained.

Description

Aircraft rapid trajectory optimization method meeting strict time and position constraints
Technical Field
The invention belongs to the field of aircraft trajectory optimization and guidance; in particular to a rapid track optimization method of an aircraft meeting strict time and position constraints.
Background
Since the 50 s of the last century, the first artificial earth satellite in the history of the launch of mankind in the soviet union, the space of human activities expanded from the atmosphere to outer space, the development of cosmos by mankind began to expand rapidly, and the development of launch vehicle technology was the basis for all space missions. In order to accurately send the effective load into a target track, the high-precision guidance technology of the carrier rocket is widely concerned by scholars at home and abroad. For years, guidance methods such as perturbation guidance, closed-circuit guidance and iterative guidance are proposed and developed rapidly, and meanwhile, an online trajectory planning algorithm represented by convex optimization is widely researched and applied to guidance instruction calculation in special states such as faults and target change. However, conventional guidance and trajectory planning methods only consider delivering the payload to the target trajectory, and do not consider payload in-orbit times and locations; for example, Songyu, autonomous trajectory planning for thrust descent fault in the ascending section of a launch vehicle, China science 2019, volume 49. Aiming at special tasks with strict time-position constraint such as geosynchronous orbit satellite launch tasks and the like, the orbit entering time and position of the effective load are not constrained in the launch vehicle guidance process, the orbit transfer of the follow-up satellite is seriously influenced, too much propellant is consumed, and the service life is reduced.
Disclosure of Invention
The invention provides an aircraft rapid trajectory optimization method meeting strict time and position constraints, which is used for solving the problems that the thrust of a high-thrust liquid rocket engine applied to a carrier rocket is not adjustable, an orbit entering point cannot be accurately controlled, and the orbit entering position cannot be constrained. Aiming at the problem that the effective load is an orbit entering task of a geosynchronous orbit satellite, the target orbit cannot be sent into by simple consideration, the geosynchronous satellite must be strictly restricted to enter the sky above a predetermined point, and otherwise, the problem that the signal interference is easily generated with other geosynchronous orbit satellites is solved.
The invention is realized by the following technical scheme:
an aircraft fast trajectory optimization method that satisfies strict time-position constraints, the aircraft fast trajectory optimization method comprising the steps of:
step 1: setting parameters; the quasi-state parameters include load at t1Time on track, standard on track point r1(ii) a Assuming that the load is t through the online re-planning and the self-adaptive guidance of the track2Time on track, actual point on track is r2；
Step 2: defining a point coordinate system; origin O of coordinate systemPIs the center of the earth, xpThe shaft is positioned on a connecting line of the geocentric and the near point of the target track and points to the near point;
and step 3: based on the parameters and point coordinate systems in the step 1 and the step 2, the concept of the approximate point angle phi is utilized to calculate the aircraft slave r1Fly to r2The required time Δ t;
and 4, step 4: and (4) correcting the satellite orbit entering time deviation by using the step 1-3 and a core secondary boot time iterative correction method.
Further, step 1 is specifically that the aircraft runs from r1Fly freely to r2The required time is Deltat, if satisfied
t1+Δt＝t2
Then explain at t1Time at r1Payload that is just about to be in orbit2Time free flight to r2. In this case, the payload may be considered to be at t2Time at r2On track, with payload at t1Time at r1And the track entering is equivalent, and the task requirement can be met.
Further, the step 2 is specifically that, if the target track is a circular track, the target track may be replaced by an intersection point because there is no near point. X is to bepThe axis being in the orbital direction omega in the target orbital planepRotated 90 DEG to obtain zpAxis perpendicular xpOpypPlane and xpAxis, ypThe axes constitute a right-hand coordinate system.
Further, the concept of using the angle phi of approach point in the step 3 is specifically that the aircraft and OXPPerpendicular to the axis, the focus of an auxiliary circle having the center of the ellipse and the semimajor axis of the ellipse as the radius, the line connecting the center of the ellipse and OXPThe included angle of the axes is phi from the flying time of the satellite in the near place as a starting point and the angle from the free flying to the near point according to the Kepler time equationpTime t ofpComprises the following steps:
Figure GDA0003149508340000021
wherein a and e are respectively semimajor axis and eccentricity of elliptical orbit, and μ is gravity constant of earth
Further, the step 4 is specifically that,
step 4.1: when the heavy carrier rocket core is shut down for the first time in the second stage, the rocket enters a preset near-earth circle transition orbit, and under the condition, the trajectory of a shutdown gliding section can be quickly predicted according to the entry parameters
Step 4.2: when the second-time starting time of the core is given, the trajectory planning can be carried out by an autonomous trajectory planning method for the thrust descent fault of the ascending section of the carrier rocket according to the initial parameters, and the deviation E of the time of the orbit is calculated according to the trajectory planning resultt；
Step 4.3: can deviate the track-in time EtConsidered as shutdown glide period time thAnd solving the deviation E meeting the orbit entry time by an iterative methodtCore two-stage boot time t of 0h。
Further, the step 4.1 is specifically,
step 4.1.1: based on core secondary rocket engine failure conditions, i.e. thrust loss coefficient kappa2Second-order core working flight segment standard time t3bAnd the time-of-flight deviation Δ t at the end of the secondary operation of the core21Can estimate the starting time of the sliding section
Figure GDA0003149508340000031
Figure GDA0003149508340000032
Step 4.1.2: set the iteration number k to 0
Step 4.1.3: according to the formula, the secondary boot time of the computing core is
Figure GDA0003149508340000033
In the initial state, an autonomous trajectory planning method for thrust descent fault of the ascending section of the carrier rocket is applied to plan the trajectory, and the flight time deviation is calculated and applied
Figure GDA0003149508340000034
Step 4.1.4: if the time of flight deviates
Figure GDA0003149508340000035
I.e. the accuracy requirement is met, thenStopping iteration to obtain the secondary starting time of the core meeting the task requirement
Figure GDA0003149508340000036
Jump to step 4.1.7, otherwise continue step 4.1.5
Step 4.1.5: make the secondary starting time of the core be
Figure GDA0003149508340000037
Calculating the time-of-flight deviation in the same way
Figure GDA0003149508340000038
Step 4.1.6: the secondary starting time of the updating core is
Figure GDA0003149508340000039
Step 4.1.7: after the calculation is finished, according to the online track planning result, the instructions of the core secondary starting time guidance pitch angle and the yaw angle are obtained
Figure GDA00031495083400000310
And adjust rocket attitude to ensure
Figure GDA00031495083400000311
And the guidance instruction meets the requirement when the computer is started.
The invention has the beneficial effects that:
the invention ensures the time position constraint precision of the on-orbit operation of the effective load and lays a foundation for the completion of subsequent tasks.
In the Geosynchronous Transfer Orbit (GTO) orbit carrying task under the thrust loss fault state, the flight time and the orbit entering position of the rocket have larger deviation, and the load is required to be difficult to enter the orbit at a preset time and a preset point.
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FIG. 1 is a flow chart of the method of the present invention.
FIG. 2 is a schematic diagram of the off-near angle φ of the present invention.
Detailed Description
The technical solutions in the embodiments of the present invention will be described clearly and completely with reference to the accompanying drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
1-2, a method for fast trajectory optimization of an aircraft satisfying strict time-position constraints, comprising the steps of:
step 1: setting parameters; the quasi-state parameters include load at t1Time on track, standard on track point r1(ii) a Assuming that the load is t through the online re-planning and the self-adaptive guidance of the track2Time on track, actual point on track is r2；
Step 2: defining a point coordinate system; origin O of coordinate systemPIs the center of the earth, xpThe shaft is positioned on a connecting line of the geocentric and the near point of the target track and points to the near point;
and step 3: based on the parameters and point coordinate systems in the step 1 and the step 2, the concept of the approximate point angle phi is utilized to calculate the aircraft slave r1Fly to r2The required time Δ t;
and 4, step 4: and (4) correcting the satellite orbit time deviation by using the step 1-3 and a core secondary boot time iterative correction method.
Further, step 1 is specifically that the aircraft runs from r1Fly freely to r2The required time is Deltat, if satisfied
t1+Δt＝t2
Then explain at t1Time at r1Payload that is just about to be in orbit2Time free flight to r2. In this case, the payload may be considered to be at t2Time at r2On track, with payload at t1Time at r1And the track entering is equivalent, and the task requirement can be met.
Further, the step 2 is specifically that, if the target track is a circular track, the target track may be replaced by an intersection point because there is no near point. X is to bepThe axis being in the orbital direction omega in the target orbital planepRotated 90 DEG to obtain zpAxis perpendicular xpOpypPlane and xpAxis, ypThe axes constitute a right-hand coordinate system.
Further, the concept of using the angle phi of approach point in the step 3 is specifically that the aircraft and OXPPerpendicular to the axis, the focus of an auxiliary circle having the center of the ellipse and the semimajor axis of the ellipse as the radius, the line connecting the center of the ellipse and OXPThe angle of the axes (as shown in fig. 2). According to the Kepler time equation, the flying time of the satellite at the near place is taken as a starting point, and the angle from the free flying to the near point is phipTime t ofpComprises the following steps:
Figure GDA0003149508340000041
wherein a and e are respectively semimajor axis and eccentricity of elliptical orbit, and μ is gravity constant of earth
Further, the step 4 is specifically that,
step 4.1: when the heavy carrier rocket core is shut down for the first time in the second stage, the rocket enters a preset near-earth circle transition orbit, and under the condition, the trajectory of a shutdown gliding section can be quickly predicted according to the entry parameters
Step 4.2: when the second-time starting time of the core is given, the trajectory planning can be carried out by an autonomous trajectory planning method for the thrust descent fault of the ascending section of the carrier rocket according to the initial parameters, and the deviation E of the time of the orbit is calculated according to the trajectory planning resultt(ii) a The autonomous trajectory planning method for the thrust descent fault of the ascending section of the carrier rocket is a known technology in the field, and is not described again;
step 4.3: can deviate the track-in time EtIs regarded as the shutdown glide period time (i.e. the secondary startup time of the core)hAs a function of (a) or (b),and solving the deviation E meeting the requirement of the time of the orbit through an iteration methodtCore two-stage boot time t of 0h
Further, the step 4.1 is specifically,
step 4.1.1: based on core secondary rocket engine failure conditions, i.e. thrust loss coefficient kappa2Second-order core working flight segment standard time t3bAnd the time-of-flight deviation Δ t at the end of the secondary operation of the core21Can estimate the starting time of the sliding section
Figure GDA0003149508340000051
Figure GDA0003149508340000052
Step 4.1.2: set the iteration number k to 0
Step 4.1.3: according to the formula, the secondary boot time of the computing core is
Figure GDA0003149508340000053
In the initial state, an autonomous trajectory planning method for thrust descent fault of the ascending section of the carrier rocket is applied to plan the trajectory, and the flight time deviation is calculated and applied
Figure GDA0003149508340000054
Step 4.1.4: if the time of flight deviates
Figure GDA0003149508340000055
Namely, the precision requirement is met, the iteration is stopped, and the secondary starting time of the core meeting the task requirement is obtained
Figure GDA0003149508340000056
Jump to step 4.1.7, otherwise continue step 4.1.5
Step 4.1.5: make the secondary starting time of the core be
Figure GDA0003149508340000057
Calculating the time-of-flight deviation in the same way
Figure GDA0003149508340000058
Step 4.1.6: the secondary starting time of the updating core is
Figure GDA0003149508340000059
Step 4.1.7: after the calculation is finished, according to the online track planning result, the instructions of the core secondary starting time guidance pitch angle and the yaw angle are obtained
Figure GDA00031495083400000510
And adjust rocket attitude to ensure
Figure GDA00031495083400000511
And the guidance instruction meets the requirement when the computer is started.

Claims (4)

1. An aircraft rapid trajectory optimization method that satisfies strict time-position constraints, the aircraft rapid trajectory optimization method comprising the steps of:
step 1: setting parameters; the parameters in the quasi-state include the load at t1Time on track, standard on track point r1(ii) a Assuming that the load is t through the online re-planning and the self-adaptive guidance of the track2Time on track, actual point on track is r2；
Step 2: defining a point coordinate system; origin O of coordinate systemPIs the center of the earth, xpThe shaft is positioned on a connecting line of the geocentric and the near point of the target track and points to the near point;
and step 3: based on the parameters and point coordinate systems in the step 1 and the step 2, the concept of the approximate point angle phi is utilized to calculate the aircraft slave r1Fly to r2Time Δ t of;
and 4, step 4: correcting the satellite orbit entering time deviation by using the step 1-3 and a core secondary boot time iterative correction method;
the step 4 is specifically that the step of,
step 4.1: after the heavy carrier rocket core is shut down for the first time in the second stage, the rocket enters a preset near-earth circle transition orbit, and the shutdown gliding section track can be quickly predicted according to the orbit entering parameters under the condition;
step 4.2: when the second-time starting time of the core is given, the trajectory planning can be carried out by an autonomous trajectory planning method for the thrust descent fault of the ascending section of the carrier rocket according to the initial parameters, and the deviation E of the time of the orbit is calculated according to the trajectory planning resultt；
Step 4.3: deviation of track entry time EtConsidered as shutdown glide period time thAnd solving the deviation E meeting the orbit entry time by an iterative methodtCore two-stage boot time t of 0h；
The step 4.1 is specifically that,
step 4.1.1: based on core secondary rocket engine failure conditions, i.e. thrust loss coefficient kappa2Second-order core working flight segment standard time t3bAnd the time-of-flight deviation Δ t at the end of the secondary operation of the core21Can estimate the starting time of the sliding section
Figure FDA0003484169400000011
Figure FDA0003484169400000012
Step 4.1.2: setting the iteration number k to be 0;
step 4.1.3: according to the formula, the secondary boot time of the computing core is
Figure FDA0003484169400000013
In the initial state, an autonomous trajectory planning method for thrust descent fault of the ascending section of the carrier rocket is applied to plan the trajectory, and the flight time deviation is calculated and applied
Figure FDA0003484169400000014
Step 4.1.4: if the time of flight deviates
Figure FDA0003484169400000015
Namely, the precision requirement is met, the iteration is stopped, and the secondary starting time of the core meeting the task requirement is obtained
Figure FDA0003484169400000021
Skipping to step 4.1.7, otherwise, continuing to step 4.1.5;
step 4.1.5: make the secondary starting time of the core be
Figure FDA0003484169400000022
Calculating the time-of-flight deviation in the same way
Figure FDA0003484169400000023
Step 4.1.6: the secondary starting time of the updating core is
Figure FDA0003484169400000024
Step 4.1.7: after the calculation is finished, according to the online track planning result, the instructions of the core secondary starting time guidance pitch angle and the yaw angle are obtained
Figure FDA0003484169400000025
ψ30And adjusting rocket attitude to ensure
Figure FDA0003484169400000026
And the guidance instruction meets the requirement when the computer is started.
2. The method for optimizing the fast trajectory of an aircraft according to claim 1, wherein the step 1 is to optimize the fast trajectory of the aircraft based on the strict time-position constraint1Fly freely to r2The required time is Deltat, if satisfied
t1+Δt＝t2
Then explain at t1Time at r1Payload that is just about to be in orbit2Time free flight to r2(ii) a In this case, the payload may be considered to be at t2Time at r2On track, with payload at t1Time at r1And the track entering is equivalent, and the task requirement can be met.
3. The method for optimizing the fast trajectory of the aircraft according to claim 1, wherein the step 2 is specifically to replace the target trajectory with a lifting intersection point if the target trajectory is a circular trajectory because there is no near point; x is to bepThe axis being in the orbital direction omega in the target orbital planepRotated 90 DEG to obtain zpAxis perpendicular xpOpypPlane and xpAxis, ypThe axes constitute a right-hand coordinate system.
4. Method for the optimization of the fast trajectory of an aircraft satisfying strict time-position constraints according to claim 1, characterized in that said concept of using the angle of approach φ of step 3 is embodied in such a way that the aircraft and OXPPerpendicular to the axis, the focus of an auxiliary circle having the center of the ellipse and the semimajor axis of the ellipse as the radius, the line connecting the center of the ellipse and OXPThe included angle of the axes is phi from the flying time of the satellite in the near place as a starting point and the angle from the free flying to the near point according to the Kepler time equationpTime t ofpComprises the following steps:
Figure FDA0003484169400000027
in the formula, a and e are respectively the semimajor axis and eccentricity of the elliptical orbit, and mu is the gravity constant of the earth.
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